FIG. 3 shows a portion of prior art reverse flow annular combustor designated by reference numeral 2. An outer transition liner 6 has a shroud 5 disposed about its outer surface. The shroud 6 abuts the liner 6 at a plurality of points 4 to define cooling air passages 3 therebetween. The liner 6 is attached to a combustor wall 8 which has a plurality of holes for injection cooling air for dilution mixing with the hot combustion gas. Dilution mixing of this cool air with the hot combustion gas immediately downstream of the flame zone is well known in the art. The dilution air is used to properly mix the hot gas, thus eliminating hot spots or streaks in the gas flow and assuring a uniform temperature profile. During combustion, the liner 6 is exposed to the hot gas exiting the combustion chamber and therefore requires cooling. This cooling is provided by a portion of the high pressure air produced by the compressor 7, represented by arrows 9, which flows through the cooling passages 3 in a radially inward, (i.e. towards the engine centerline), direction exiting as low momentum air at the inner portion of the liner 6 and is then dumped into the gas stream upstream of the first stage turbine stator, not shown.
The amount of air flow through the cooling passages is a function of the pressure drop from the inlet of the cooling passages to their exit. The greater the pressure drop the larger the cooling flow. In the prior art, this pressure drop has been limited by two factors. First, prior art cooling passages only extend to just upstream of the first stage stator, and second, the first stage stators generate horseshoe vortices at their leading edges which produce local regions of increased pressure. Accordingly, there is need in gas turbine engines for a combustor-to-turbine transition assembly that overcomes the prior art limitations.